Supersonic aircraft are typically equipped with gas turbine engines to achieve supersonic speeds. Gas turbine engines include compressors that require subsonic flow in a range of about Mach 0.3 to 0.6 at the face of the engine. An inlet decelerates the incoming supersonic air flow to a speed compatible with the requirements of the gas turbine engine. In conventional propulsion systems, supersonic inlets include a compression surface and corresponding flow path that decelerates the supersonic flow using a strong normal shock. Downstream of the normal shock, subsonic flow is further decelerated using a diffuser to a speed corresponding with requirements of the gas turbine engine. This normal shock places a sharp pressure rise on the subsonic boundary layer which causes the boundary layer to thicken and separate. This phenomenon results in lower inlet performance and higher distortion to the engine.
Traditional inlet design methods have generally focused on improving propulsion system performance by maximizing total inlet pressure recovery and hence gross engine thrust. While high pressure recovery definitely provides certain gains, maximizing pressure recovery typically comes at the price of significant inlet drag and inlet complexity, characteristics that typically run counter to a robust and low cost-of-operation design. For example, attempts to increase pressure recovery include bleed air-based methods, which improve inlet pressure recovery through shock strength management and boundary layer removal. However, bleed air-based methods typically take a large portion of the intake flow to produce the desired results and suffer corresponding drag-related penalties once the bleed flow is eventually exhausted overboard. Additionally, extensive secondary systems are typically required, consisting of complex flow routing equipment. The stabilizing bleed system represents an additional loss in net thrust of the system, as it requires added pressure loss (or mechanical pumping) to induce the bleed flow.